It is known that, at the rear of a bypass turbojet engine mounted on an aircraft, the supersonic cold flow, flowing in the downstream direction of said turbojet engine, comes into contact with the exterior aerodynamic airstream of said turbojet engine. Since the speeds of said cold flow and of said airstream differ from one another, this results in inter-penetrating fluidic shear, which generates noise, known in aeronautical parlance as “jet noise”.
In addition, as a result of a discontinuity in static pressure between the external pressure and the pressure at the exit from the jet pipe, this supersonic cold flow gives rise to a series of compression-expansion (speed fluctuation) cells which act as noise amplifiers and produce a noise known in aeronautical parlance as “shock cell noise”, this English-language term “shock cell noise” being widely recognized.
In order to attenuate the noise emitted at the rear of a bypass turbojet engine, the idea of modifying the rear part of the cold flow nozzle has already been actioned. For example, extending said nozzle rearward using “chevrons” (see, for example, U.S. Pat. Nos. 4,284,170 and 6,360,528) or shaping the rear part of said nozzle in the form of “undulating lobes” (see, for example, GB 2 160 265, U.S. Pat. Nos. 4,786,016 and 6,082,635) have already been proposed.
Aside from the fact that these known nozzles demand definitive special shapes which in general increase the cost, mass and drag, it should be pointed out that, although they are effective in attenuating jet noise by creating turbulence that encourages the cold flow and the exterior aerodynamic airstream to mix, they have only a very limited effect in reducing shock cell noise.
Another source, document EP-1 703 114, describes a reduced-noise turbojet engine in which a plurality of bosses are distributed at the periphery of the outlet orifice of the cold flow, projecting into the latter, each of said bosses forming a convergent followed by a divergent connected to the edge of said cold flow outlet orifice.